U.S. patent application number 13/326654 was filed with the patent office on 2012-04-26 for spacecraft momentum management.
This patent application is currently assigned to SPACE SYSTEMS/LORAL, INC.. Invention is credited to Kam K. Chan, Mohammad Saghir Munir, Byoungsam Woo.
Application Number | 20120097797 13/326654 |
Document ID | / |
Family ID | 45972143 |
Filed Date | 2012-04-26 |
United States Patent
Application |
20120097797 |
Kind Code |
A1 |
Woo; Byoungsam ; et
al. |
April 26, 2012 |
SPACECRAFT MOMENTUM MANAGEMENT
Abstract
Three-axis spacecraft momentum management is performed for a
spacecraft traveling along a trajectory, by an actuator including
at least one thruster disposed on a single positioning mechanism.
As the spacecraft travels along the trajectory, a desired line of
thrust undergoes a substantial rotation in inertial space. When the
spacecraft is located at a first location on the trajectory, the
single positioning mechanism orients the thruster so as to produce
a first torque to manage stored momentum in at least one of a first
and a second of the three inertial spacecraft axes. When the
spacecraft is located at a second location on the trajectory, the
single positioning mechanism orients the thruster so as to produce
a second torque to manage stored momentum in at least a third of
the three inertial spacecraft axes.
Inventors: |
Woo; Byoungsam; (San Jose,
CA) ; Munir; Mohammad Saghir; (Union City, CA)
; Chan; Kam K.; (Saratoga, CA) |
Assignee: |
SPACE SYSTEMS/LORAL, INC.
Palo Alto
CA
|
Family ID: |
45972143 |
Appl. No.: |
13/326654 |
Filed: |
December 15, 2011 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
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12925386 |
Oct 20, 2010 |
|
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13326654 |
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Current U.S.
Class: |
244/158.6 ;
701/3 |
Current CPC
Class: |
B64G 1/26 20130101; B64G
1/242 20130101; F03H 1/0037 20130101 |
Class at
Publication: |
244/158.6 ;
701/3 |
International
Class: |
B64G 1/10 20060101
B64G001/10; G05D 1/00 20060101 G05D001/00; B64G 1/24 20060101
B64G001/24 |
Claims
1. A method for managing momentum of a spacecraft traveling along a
trajectory, the method comprising: determining a respective
momentum storage error (MSE) in each of three inertial spacecraft
axes, said respective MSE comprising a difference, for each axis,
between a momentum value actually stored on the spacecraft and a
desired momentum value; reducing each respective MSE by producing,
with at least one thruster disposed on a single positioning
mechanism, a plurality of torques, by: orienting the thruster, with
the single positioning mechanism, so as to produce a first torque
that reduces the respective MSE of either or both of a first and a
second of the three inertial spacecraft axes when the spacecraft is
located at a first location on the trajectory, and, orienting the
thruster, with the single positioning mechanism, so as to produce a
second torque that reduces the respective MSE of at least a third
of the three inertial spacecraft axes when the spacecraft is
located at a second location on the trajectory; wherein, the single
positioning mechanism is configured to orient the thruster so as to
simultaneously (i) accelerate the spacecraft along a line of thrust
and (ii) produce a torque around at least one of two axes
substantially orthogonal to the nominal thrust vector; and as the
spacecraft travels along the trajectory, a desired line of thrust
undergoes a substantial rotation in inertial space.
2. The method as recited in claim 1, wherein spacecraft
acceleration and MSE are simultaneously controlled by the at least
one thruster disposed on the single positioning mechanism.
3. The method as recited in claim 1, wherein the thruster is a low
thrust electric propulsion device.
4. The method as recited in claim 3, wherein the thruster is a Hall
effect thruster.
5. The method as recited in claim 1, wherein, as the spacecraft
travels along the trajectory, the substantial rotation is
approximately ninety degrees.
6. The method as recited in claim 1, wherein the single positioning
mechanism has two degrees of freedom.
7. A spacecraft comprising: at least one thruster; spacecraft
control electronics configured to: (i) generate a desired orbit
transfer profile for the spacecraft; and (ii) determine a
respective momentum storage error (MSE) in each of three inertial
spacecraft axes, said respective MSE comprising a difference, for
each axis, between a momentum value actually stored on the
spacecraft and a desired momentum value; and a spacecraft steering
apparatus, comprising the at least one thruster disposed on a
single positioning mechanism, that, responsive to signals from the
spacecraft control electronics: controls the attitude of the
satellite so as to follow the desired orbit transfer profile; and
reduces each respective MSE by producing, with the at least one
thruster, a plurality of torques, by: orienting the thruster, with
the single positioning mechanism, so as to produce a first torque
that reduces the respective MSE of either or both of a first and a
second of the three inertial spacecraft axes when the spacecraft is
located at a first location on the trajectory, and, orienting the
thruster, with the single positioning mechanism, so as to produce a
second torque that reduces the respective MSE of at least a third
of the three inertial spacecraft axes when the spacecraft is
located at a second location on the trajectory; wherein, the single
positioning mechanism is configured to orient the thruster so as to
simultaneously (i) accelerate the spacecraft along a line of thrust
and (ii) produce a torque around at least one of two axes
substantially orthogonal to the nominal thrust vector; and as the
spacecraft travels along the trajectory, a desired line of thrust
undergoes a substantial rotation in inertial space.
8. The spacecraft as recited in claim 7 wherein the at least one
thruster comprises an electric propulsion thrusters
9. The spacecraft as recited in claim 8 wherein the at least one
thruster comprises a Hall effect thruster.
10. The spacecraft as recited in claim 7, wherein, as the
spacecraft travels along the trajectory, the substantial rotation
is approximately ninety degrees.
11. The spacecraft as recited in claim 7 wherein the spacecraft
control electronics comprises a profile generator configured to
compute a desired orbit transfer profile such that perigee, apogee
and inclination of the spacecraft are adjusted simultaneously in a
mass-efficient manner.
12. The spacecraft as recited in claim 7 wherein the desired orbit
transfer profile includes: placing the spacecraft in an
Earth-pointed attitude when the spacecraft is at a predefined point
in the trajectory; slewing the spacecraft from the Earth-pointed
attitude to a desired orbit raising attitude; and steering the
spacecraft according to the desired orbit transfer profile while
changing a spacecraft velocity.
13. The method as recited in claim 7, wherein the single
positioning mechanism has two degrees of freedom.
14. A method comprising: dynamically computing an optimal steering
profile for a spacecraft, based on position of the spacecraft on a
trajectory, the spacecraft comprising at least one thruster
disposed on a single positioning mechanism and an inertial
reference sensor; dynamically computing the spacecraft's actual
position; steering the spacecraft according to the computed optimal
steering profile such that the at least one thruster imparts a
change in velocity of the spacecraft along a desired direction;
periodically shutting down the at least one thruster and
reorienting the spacecraft; restarting the at least one thruster;
autonomously repeating the above steps until the desired orbit is
reached wherein three axis momentum management of the spacecraft is
performed by: determining a respective momentum storage error (MSE)
in each of three inertial spacecraft axes, said respective MSE
comprising a difference, for each axis, between a momentum value
actually stored on the spacecraft and a desired momentum value;
reducing each respective MSE by producing, with at least one
thruster disposed on a single positioning mechanism, a plurality of
torques, by: orienting the thruster, with the single positioning
mechanism, so as to produce a first torque that reduces the
respective MSE of either or both of a first and a second of the
three inertial spacecraft axes when the spacecraft is located at a
first location on the trajectory, and, orienting the thruster, with
the single positioning mechanism, so as to produce a second torque
that reduces the respective MSE of at least a third of the three
inertial spacecraft axes when the spacecraft is located at a second
location on the trajectory; wherein, the single positioning
mechanism is configured to orient the thruster so as to
simultaneously (i) accelerate the spacecraft along a line of thrust
and (ii) produce a torque around at least one of two axes
substantially orthogonal to the nominal thrust vector; and as the
spacecraft travels along the trajectory, a desired line of thrust
undergoes a substantial rotation in inertial space.
15. The method as recited in claim 14, wherein the at least one
thruster comprises a Hall effect thruster.
16. The method as recited in claim 14, where the inertial
references sensor comprises a gyro that is reset to remove any
drift when the spacecraft is in an Earth pointed orientation, using
a calculated position of the Earth relative to the spacecraft,
spacecraft orbital information and Earth sensor data.
17. The method as recited in claim 14, wherein the single
positioning mechanism has two degrees of freedom.
Description
RELATED APPLICATION
[0001] This application is a continuation-in-part of U.S.
application Ser. No. 12/925,386, filed Oct. 20, 2010, the
disclosure of which is hereby incorporated by reference in its
entirety for all purposes.
TECHNICAL FIELD
[0002] This invention relates generally to spacecraft momentum
management and, in particular, to providing three axes momentum
management during orbit transfer maneuvers such as orbit
raising.
BACKGROUND
[0003] The assignee of the present invention manufactures and
deploys spacecraft for, commercial, defense and scientific
missions. On board propulsion systems of such spacecraft are
frequently required to perform orbit raising (or transfer). For
example, there is frequently a requirement for commercial
spacecraft to perform orbit raising from a launch vehicle transfer
orbit to a geosynchronous orbit. As a further example, certain
missions may require transfers between orbits. Such maneuvers may
be performed with chemical thrusters, or with one or more with low
thrust electric thrusters, as described by Oh, U.S. Pat. No.
6,543,723 (hereinafter "Oh"), assigned to the assignee of the
present invention, and Gelon, et al., U.S. Pat. No. 7,113,851,
(hereinafter "Gelon") entitled "Practical Orbit Raising System and
Method for Geosynchronous Satellites" assigned to the assignee of
the present invention, and Gelon.
[0004] Known orbit raising techniques are also described in U.S.
Pat. No. 5,595,360 issued to Spitzer, entitled "Optimal Transfer
Orbit Trajectory Using Electric Propulsion," U.S. Pat. No.
6,116,543, issued to Koppel, entitled "Method and a System for
Putting a Space Vehicle into Orbit, Using Thrusters of High
Specific Impulse."
[0005] Characteristically, during such transfers, spacecraft
momentum has to be managed so as to provide three axis attitude
control. Momentum storage systems are employed to store accumulated
momentum resulting from a disturbance torque environment, and
thereby reduce the pointing disturbance and propellant usage
associated with a thruster actuation. These systems, consisting of
reaction wheels, have a storage capacity that may be described in
terms of a permissible range of wheel speeds. As a result, a
momentum management strategy must use thrusters or other actuators
such as magnetic torquers or solar sailing techniques to unload
momentum in order to prevent wheel speeds from going outside the
permissible range.
[0006] Known orbit raising techniques provide momentum management
during long duration operation of electric propulsion thrusters by
gimbaling and/or throttling the thruster(s) providing the orbit
raising velocity change. Where, as is desirable for reliability and
cost reasons, orbit raising is to be performed with thruster(s)
mounted on a single positioning mechanism, a problem arises that
such a single gimbaled thruster can only provide torque about the
two axes orthogonal to its thrust axis. Thus, it is not possible to
generate torque parallel to the thrust vector. Conventionally, this
problem is solved by providing at least one additional actuator to
provide yaw authority.
[0007] As a result, system performance is penalized by the
additional hardware cost, mass, and complexity.
SUMMARY
[0008] The present inventors have recognized that, for a spacecraft
traveling along a trajectory, three-axis spacecraft momentum
management may be advantageously performed by an actuator
consisting of as few as one thruster disposed on a single
positioning mechanism. As the spacecraft travels along the
trajectory, a desired line of thrust undergoes a substantial
rotation in inertial space, and the single positioning mechanism is
configured to orient the thruster so as to simultaneously (i)
accelerate the spacecraft along a line of thrust and (ii) produce a
torque around at least one of two axes substantially orthogonal to
the nominal thrust vector. When the spacecraft is located at a
first location on the trajectory, the single positioning mechanism
orients the thruster so as to produce a first torque to manage
stored momentum in at least one of a first and a second of the
three inertial spacecraft axes. When the spacecraft is located at a
second location on the trajectory, the single positioning mechanism
orients the thruster so as to produce a second torque to manage
stored momentum in at least a third of the three inertial
spacecraft axes.
[0009] In an embodiment, momentum of a spacecraft traveling along a
trajectory is managed by determining a respective momentum storage
error (MSE) in each of three inertial spacecraft axes. The
respective MSE is a difference, for each axis, between a momentum
value actually stored on the spacecraft and a desired momentum
value. Each respective MSE is reduced by producing, with at least
one thruster disposed on a single positioning mechanism, a
plurality of torques, by orienting the thruster, with the single
positioning mechanism, so as to produce a first torque that reduces
the respective MSE of a first and/or a second of the three inertial
spacecraft axes when the spacecraft is located at a first location
on the trajectory. The thruster is oriented with the single
positioning mechanism, so as to produce a second torque that
reduces the respective MSE of at least a third of the three
inertial spacecraft axes when the spacecraft is located at a second
location on the trajectory. The single positioning mechanism is
configured to orient the thruster so as to simultaneously (i)
accelerate the spacecraft along a line of thrust and (ii) produce a
torque around at least one of two axes substantially orthogonal to
the nominal thrust vector. As the spacecraft travels along the
trajectory, a desired line of thrust undergoes a substantial
rotation in inertial space.
[0010] In another embodiment, spacecraft acceleration and MSE may
be simultaneously controlled by the at least one thruster disposed
on the single positioning mechanism.
[0011] In an embodiment, the thruster may be a low thrust electric
propulsion device. The thruster may be a Hall effect thruster.
[0012] In a further embodiment, as the spacecraft may travel along
the trajectory, the substantial rotation may be approximately
ninety degrees.
[0013] In another embodiment, the single positioning mechanism may
have two degrees of freedom.
[0014] In an embodiment, a spacecraft has at least one thruster and
spacecraft control electronics configured to: (i) generate a
desired orbit transfer profile for the spacecraft; and (ii)
determine a respective momentum storage error (MSE) in each of
three inertial spacecraft axes. The respective MSE is a difference,
for each axis, between a momentum value actually stored on the
spacecraft and a desired momentum value. The spacecraft also has a
spacecraft steering apparatus, including the at least one thruster
disposed on a single positioning mechanism, that, responsive to
signals from the spacecraft control electronics: controls the
attitude of the spacecraft so as to follow the desired orbit
transfer profile; and reduces each respective MSE by producing,
with the at least one thruster, a plurality of torques, by (i)
orienting the thruster, with the single positioning mechanism, so
as to produce a first torque that reduces the respective MSE of a
first and/or a second of the three inertial spacecraft axes when
the spacecraft is located at a first location on the trajectory,
and, (ii) orienting the thruster, with the single positioning
mechanism, so as to produce a second torque that reduces the
respective MSE of at least a third of the three inertial spacecraft
axes when the spacecraft is located at a second location on the
trajectory. The single positioning mechanism is configured to
orient the thruster so as to simultaneously (i) accelerate the
spacecraft along a line of thrust and (ii) produce a torque around
at least one of two axes substantially orthogonal to the nominal
thrust vector. As the spacecraft travels along the trajectory, a
desired line of thrust undergoes a substantial rotation in inertial
space.
[0015] In another embodiment, the spacecraft control electronics
may include profile generator configured to compute a desired orbit
transfer profile such that perigee, apogee and inclination of the
spacecraft are adjusted simultaneously in a mass-efficient
manner.
[0016] In an embodiment the desired orbit transfer profile may
include placing the spacecraft in an Earth-pointed attitude when
the spacecraft is at a predefined point in the trajectory; slewing
the spacecraft from the Earth-pointed attitude to a desired orbit
raising attitude; and steering the spacecraft according to the
desired orbit transfer profile while changing a spacecraft
velocity.
[0017] In an embodiment, an optimal steering profile for a
spacecraft is dynamically computed, based on position of the
spacecraft on a trajectory, the spacecraft including at least one
thruster disposed on a single positioning mechanism and an inertial
is dynamically computed. The spacecraft is steered according to the
computed optimal steering profile such that the at least one
thruster imparts a change in velocity of the spacecraft along a
desired direction. The at least one thruster is periodically shut
down and the spacecraft is reoriented. The at least one thruster is
restarted. The above steps are repeated until the desired orbit is
reached. Three axis momentum management of the spacecraft is
performed by determining a respective momentum storage error (MSE)
in each of three inertial spacecraft axes, said respective MSE
including a difference, for each axis, between a momentum value
actually stored on the spacecraft and a desired momentum value;
reducing each respective MSE by producing, with at least one
thruster disposed on a single positioning mechanism, a plurality of
torques, by: orienting the thruster, with the single positioning
mechanism, so as to produce a first torque that reduces the
respective MSE of either or both of a first and a second of the
three inertial spacecraft axes when the spacecraft is located at a
first location on the trajectory, and, orienting the thruster, with
the single positioning mechanism, so as to produce a second torque
that reduces the respective MSE of at least a third of the three
inertial spacecraft axes when the spacecraft is located at a second
location on the trajectory. The single positioning mechanism is
configured to orient the thruster so as to simultaneously (i)
accelerate the spacecraft along a line of thrust and (ii) produce a
torque around at least one of two axes substantially orthogonal to
the nominal thrust vector; and as the spacecraft travels along the
trajectory, a desired line of thrust undergoes a substantial
rotation in inertial space.
[0018] In an embodiment, the inertial references sensor includes a
gyro that is reset to remove any drift when the spacecraft is in an
Earth pointed orientation, using a calculated position of the Earth
relative to the spacecraft, spacecraft orbital information and
Earth sensor data.
BRIEF DESCRIPTION OF THE DRAWINGS
[0019] Features of the invention are more fully disclosed in the
following detailed description of the preferred embodiments,
reference being had to the accompanying drawings, in which like
reference numerals designate like structural element, and in
which:
[0020] FIG. 1 illustrates an example of an architecture of
apparatus that provides for electric propulsion satellite orbit
raising;
[0021] FIG. 2 illustrates details of an example electric propulsion
satellite orbit raising state machine;
[0022] FIG. 3a illustrates an example electric propulsion satellite
orbit raising timeline for two SPT electric orbit raising
[0023] FIG. 3b illustrates an example electric propulsion satellite
orbit raising timeline for a single SPT electric orbit raising
[0024] FIG. 4 illustrates an example electric propulsion satellite
orbit raising method.
[0025] FIG. 5a illustrates an example orientation of the dual axes
positioning mechanisms (DAPMs) for two SPT electric orbit
raising.
[0026] FIG. 5b illustrates an example orientation of the dual axes
positioning mechanism (DAPMs) for single SPT electric orbit
raising.
[0027] FIG. 6 illustrates an example orientation of the ideal
electric orbit raising thrust vector profile in inertial space.
[0028] FIG. 7 illustrates an example orientation of the Sun with
respect to the satellite during single SPT electric orbit
raising.
[0029] FIG. 8 illustrates an example momentum management method
[0030] FIG. 9 illustrates an example orbit transfer method
[0031] Throughout the drawings, the same reference numerals and
characters, unless otherwise stated, are used to denote like
features, elements, components, or portions of the illustrated
embodiments. Moreover, while the subject invention will now be
described in detail with reference to the drawings, the description
is done in connection with the illustrative embodiments. It is
intended that changes and modifications can be made to the
described embodiments without departing from the true scope and
spirit of the subject invention as defined by the appended
claims.
DETAILED DESCRIPTION
[0032] Specific exemplary embodiments of the invention will now be
described with reference to the accompanying drawings. This
invention may, however, be embodied in many different forms, and
should not be construed as limited to the embodiments set forth
herein. Rather, these embodiments are provided so that this
disclosure will be thorough and complete, and will fully convey the
scope of the invention to those skilled in the art.
[0033] The terms "spacecraft", "satellite" and "vehicle" may be
used interchangeably herein, and generally refer to any orbiting
satellite or spacecraft system.
[0034] Referring to the drawing figures, disclosed are apparatus 10
(FIG. 1) and methods 40 (FIG. 4) for raising the orbit of a
satellite using electric propulsion. In some embodiments, a state
machine, an Earth sensor, and a gyro are employed. For example,
FIG. 1 illustrates architecture of apparatus 10 that determines an
attitude profile that the satellite should preferably follow during
electric propulsion orbit raising. FIG. 2 illustrates details of an
exemplary electric propulsion satellite orbit raising state machine
14 that may be employed in the apparatus 10 shown in FIG. 1. FIG. 3
illustrates an exemplary timeline that implements electric
propulsion satellite orbit raising.
[0035] Referring to FIG. 1, satellite 11 may be configured to have
an onboard positioning system 12, such as an orbit propagator or
global positioning system (GPS). Onboard positioning system 12 may
be coupled to an electric orbit raising (EOR) profile generator 13.
EOR profile generator 13 may be coupled to EOR state machine 14.
Outputs of the EOR state machine 14, along with data outputs from
Earth sensor 16 and gyro 17 may be coupled to satellite steering
apparatus 15. The satellite steering apparatus 15 outputs error
signals that are input to control laws 18 that control the attitude
of the satellite 11 via actuators 19, such as a wheel system, for
example. Satellite 11 may be configured to have one or more
electric propulsion thrusters (not illustrated), such as Hall
effect thrusters (also referred to herein as stationary plasma
thrusters, or SPT's), and/or bi-propellant thrusters (if
desired).
[0036] In operation, and also referring to FIGS. 2 and 3, onboard
positioning system 12 is used to compute the position of the
satellite 11 in a dynamically changing orbit, accounting for
firings of the electric propulsion thrusters and (if any)
bi-propellant thrusters. The onboard positioning system 12 may
generate instantaneous orbital data that are fed into the EOR
profile generator 13. EOR profile generator 13 computes an ideal
EOR attitude (profile) that the satellite 11 must follow so that
the perigee, apogee and inclination of the satellite 11 can be
adjusted simultaneously in a mass-efficient manner. This ideal
profile is fed into the EOR state machine 14 which generates
(computes) a desired profile to steer the satellite 11 according to
the ideal profile during orbit raising (i.e., when in phase D).
Otherwise the profile generated by the EOR state machine 14 keeps
the satellite 11 Earth-pointed (i.e., when in phase A). In between,
the profile generated by the state machine 14 either slews the
satellite 11 from Earth-pointed configuration to the EOR desired
attitude (i.e., phase B), or from an EOR-desired attitude to an
Earth-pointed configuration (i.e., phase F). During phases C and E,
the electric propulsion thrusters are started and shut down by
state machine 14. Configuration of the on-board fault detection,
isolation and recovery may also be carried out by the state machine
14 at the appropriate true-anomalies provided by the onboard
positioning system 12 (orbit propagator 12 or global positing
system 12). The true anomaly is the angle measured in the direction
of motion from perigee to the position of the satellite 11 at some
defined epoch time. The EOR state machine 14 also includes phase G
which aborts any of phases B, C or D, which may be required to shut
down the electric propulsion thrusters and return the satellite 11
to an Earth pointed orientation in the event of an anomaly. The
Earth pointed orientation is generally trusted to be a safe
attitude, as it guarantees telemetry and commanding when there is
line of site coverage to a ground station.
[0037] The EOR profile generator 13 and the state machine 14 are
parameterized by the instantaneous orbital elements (i.e., the
orbital data from the positioning system 12). Therefore, given the
on-board knowledge of the orbit, the desired time varying optimal
steering attitude profile required for EOR, and the time at which
state machine 14 transitions between major phases adjusts
automatically. By the very nature of automatic Earth
reacquisitions, to upright the satellite 11 and reset the gyro 16
based on the calculated position of the Earth relative to the
satellite 11, satellite orbital knowledge and Earth sensor data
(which indirectly localizes the satellite 11 with the correct Earth
geometry), the effects of gyro drift are removed at the beginning
of every revolution. All a user has to do is monitor each
revolution of the satellite 11.
[0038] If the satellite 11 includes a star tracker, Earth
acquisition would not be required, but the state machine 14 can
still drive the entire EOR process and continue to provide the
operational ease and autonomy for which it was designed. If there
is an onboard GPS 12, then the orbit propagator 12 is also not
required, as the GPS 12 can provide the desired information.
[0039] The apparatus 10 and methods 40 simultaneously drive the
satellite perigee, apogee and inclination toward target values,
with the entire process automated by the state machine 14 driven by
the onboard positioning system 12 (orbit propagator 12 or GPS 12).
The EOR process is simplified from an operations point of view, as
a result of using the state machine 14. The state machine 14, like
the optimal steering profiles, is parameterized by the onboard
orbital data derived from the positioning system 12. The
positioning system 12 (orbit propagator 12 or GPS) accounts for the
changing orbit due to the firing of the SPTs with the satellite 11
in the optimal steering attitude. Thus, the optimal steering
profile and state machine 14 are autonomously adjusted in terms of
timing due to a dynamically changing orbit. All a ground station
operator has to do is monitor the satellite 11.
[0040] The following presents details of the EOR process
implemented by the apparatus 10 and methods 40. FIG. 2 illustrates
details of an exemplary state machine 14. FIG. 3a illustrates an
exemplary timeline for achieving electric propulsion satellite
orbit raising.
[0041] The EOR process begins with the satellite 11 in an Earth
pointed configuration, using the Earth sensor 16 (Phase A). At this
time, the satellite 11 is steered in yaw, such that the roll rate
reported by the gyro 16 is zero. Assuming that the gyro 16 is
well-calibrated, the x-axis of the satellite 11 is in the orbital
plane, with the z-axis of the satellite 11 locked onto the Earth.
This phase of the EOR state machine 14 is shown in FIGS. 2 and 3,
and is identified as phase A, and is referred to as the Earth
pointed phase.
[0042] Once the desired true anomaly is reached, the state machine
14 transitions to phase B where the gyro 15 is initialized with
respect to an inertial reference frame, assuming that the satellite
11 is Earth-pointed (satellite z-axis is Earth-pointed), and the
satellite x-axis is in the orbital plane close to the velocity
vector. From the satellite's orbit, Earth sensor data, and the
Earth's calculated position, the attitude of the satellite 11 in
inertial space can be exactly localized. Thus, an attitude
quaternion to which the gyro 17 must be initialized can be
computed. Immediately after initialization of the gyro 17, the
satellite 11 executes a large slew using only the gyro 17 for
inertial reference and reaction wheels to align the thrust vector
of the satellite 11 with the desired EOR profile.
[0043] Once on the profile, the EOR state machine 14 is
transitioned to phase C where the electric propulsion thrusters are
ignited, and the EOR state machine 14 then transitions to phase, D.
For the next 18-20 hours, while in phase D, the satellite 11 is
steered according to the EOR profile while in reaction-wheel-based
gyro mode. During this time the momentum is managed by offsetting
the combined thrust vector from the center of mass of the satellite
11.
[0044] Once the desired true-anomaly is reached, the electric
propulsion thrusters are turned off in phase E, and then the
satellite 11 is slewed back toward the Earth using the
reaction-wheel-based gyro mode, with knowledge of the Earth
provided by the positioning system 12 (orbit propagator 12 or GPS
12). Once the Earth is acquired, there is likely to be a small
offset reported in the Earth measurement (from the Earth sensor
16), due to gyro drift and errors in the time and onboard orbital
data. However a majority of this error (>95%) should be due to
gyro drift alone. Once control switches to the Earth sensor 16, and
the satellite 11 locks onto the Earth, the error due to the gyro 17
is removed upon next reinitialization of the onboard
gyro-propagated attitude estimate in phase A. The cycle is then
repeated.
[0045] It should be evident that if a user re-initializes the orbit
before the onboard orbital error exceeds some predetermined amount,
the entire process remains well automated. Alternatively having an
onboard GPS 12 can do the same thing.
[0046] With the above in mind, FIG. 4 illustrates an exemplary
electric propulsion satellite orbit raising method 40. The
exemplary method 40 dynamically computes 41 the position of the
satellite 11 onboard the satellite 11. An optimal satellite
steering profile is also dynamically computed 42 onboard the
satellite 11, based on the satellite's position in orbit. This
optimal profile provides more mass benefit than prior art solutions
because it simultaneously drives the satellite perigee, apogee and
inclination toward target values.
[0047] The satellite 11 is steered 42, using the gyro 17 (or star
tracker) for inertial reference, according to the computed profile
such that a change in velocity (delta-V) of the satellite 11 is
imparted in a desired direction.
[0048] Periodically (as often as needed), the propulsion system is
shut down and the Earth is reacquired 43 via direct slew in
conjunction with an onboard orbit propagator, for example, to
re-initialize 44 the gyro and perform other satellite maintenance,
if desired, and then the propulsion system is restarted 45 once on
the profile again. This removes the effect of accumulated gyro
drift. The reacquisition in order to reset the gyro, maybe needed
as often as every revolution or as infrequently as every few days.
In the case of a star tracker, periodic Earth re-acquisition 43 is
not required.
[0049] Steps 41 through 45 are autonomously repeated 46 until the
desired orbit is reached. No daily planning to compute the EOR
attitude profile is required, as the profile is computed onboard
the satellite 11 using orbital data from the onboard positioning
system 12 (orbit propagator 12 or global positing system 12).
[0050] The method 40 described herein, which may last on the order
of several months, automates the entire EOR process, using the
state machine 14 which issues commands parameterized by the
true-anomaly of the dynamically changing orbit. Since the
expressions for the optimal profiles are also parameterized by the
orbital data, the use of the on-board positioning system 12 (orbit
propagator 12 or global positing system 12) that accounts for the
imparted change in velocity (delta-V) is made. This ensures that
the on-board dynamic positioning system 12 (orbit propagator 12 or
global positing system 12) is properly initialized, automates the
entire process, including autonomously reacquiring the Earth to
reset the gyro 17 in order to remove the accumulated drift. This
also eliminates the requirement for star trackers, which amounts
for increased cost and mass.
[0051] Furthermore, using a state machine 14 reduces the chances of
operator error and allows the satellite 11 to fly through telemetry
and command outages, and streamlines the entire process. The
automation provided by the state machine 14 requires that the
satellite 11 only have brief coverage to just a single ground
station for periodic maintenance, thus reducing orbit raising costs
compared with prior art solutions.
[0052] The illustrative drawings showing the EOR timeline (FIG. 3a
and FIG. 3b) depict exemplary profiles, where the satellite
thruster vector is primarily steered in the orbital plane. However
it is important to mention that the present invention also applies
when the satellite 11 is steered out of plane, in the case of
inclination removal or adjustment.
[0053] When EOR is performed with two SPTs mounted on separate dual
axes positioning mechanisms (DAPMs) as depicted in FIG. 5a, full
3-axis momentum dumping is possible using techniques well known in
the art. For example, in the illustration depicted in FIG. 5a, roll
and pitch momentum may be dumped by actuating the north and south
DAPMs together in the same direction, while yaw momentum may be
dumped by actuating the north and south DAPMs in equal opposite
directions. Thus the use of chemical thrusters for momentum dumping
with SPTs mounted on two DAPMs is not required.
[0054] In an alternative embodiment, EOR may be performed with as
little as a single thruster disposed on a single positioning
mechanism. Referring now to FIG. 5b, thruster 510 may be aimed by
positioning mechanism 520 through the center of mass of satellite
11 such that the thrust vector (or "line of thrust") is aligned
with the desired line of thrust.
[0055] FIG. 3b illustrates an exemplary timeline for achieving
electric propulsion satellite orbit raising with a single SPT fired
through the satellite center of mass, but kept on the same EOR
profile as that shown in FIG. 3a.
[0056] When EOR is performed with only one SPT, or with more than
one SPT mounted on the same DAPM, momentum can only be dumped in
two of the three axes at any moment of time. This is because
offsetting the net thrust vector from the center of mass produces
only torques orthogonal to the thruster vector (or "line of
thrust"). At any moment of time, it is not possible to generate
torque parallel to the thrust vector, hence the direction in which
momentum cannot be dumped is the same as the thrust vector itself.
From this, one can incorrectly assume that use of additional
actuators may be required to manage momentum in the direction of
the thrust vector. However, this is not the case, because the line
of thrust (i.e. the unmanageable axis) rotates in inertial
space.
[0057] For example, referring now to FIG. 6, a spacecraft
trajectory is illustrated as viewed from an angle orthogonal to the
orbital plane of the trajectory. The desired line of thrust at each
of a series of locations along the trajectory is depicted by a
vector arrow. For example, vector arrow 601(i) and vector arrow
601(n) illustrate the desired line of thrust near respective
locations 60 and 61. It may be observed that, between location 60
and 61, the desired line of thrust undergoes a substantial rotation
in inertial space. In the illustrated example, the rotation is
approximately ninety degrees. This net rotation may be provided by
steering satellite 11 to which the thruster is mounted.
[0058] The present inventors have appreciated that, using the
presently disclosed techniques, the momentum wheels may accumulate
momentum in an initially unmanageable axis, until the satellite 11
is steered to an orientation where that momentum can be dumped.
Therefore, even though at any moment of time, only two axes of
momentum can be dumped, the net rotation of the desired line of
thrust in inertial space implies that, within a determinable period
of time, the manageable axes may rotate by 90 degrees, thereby
spanning the previously unmanageable portion of the 3D space. Thus,
given adequately sized momentum wheels, and a desired line of
thrust that undergoes a substantial rotation in inertial space,
three axis momentum management may be provided by a thruster
disposed on a single positioning mechanism.
[0059] Although a detailed explanation of an implementation using
SPTs for electric orbit raising has been described above, the
present momentum management techniques may be employed for many
types of orbit transfer missions, and with chemical or electric
thrusters of many types, provided only that the desired line of
thrust makes a substantial rotation in inertial space within a
period of time commensurate with the momentum storage capability of
the spacecraft. The above mentioned constraint may be met for many
spacecraft missions that require attitude change in inertial space.
Such missions may include central body (e.g., Earth, Moon or Sun)
pointing missions as well as continuous delta V missions such as
low thrust orbit raising missions. In such missions, the attitude
of the spacecraft ordinarily undergoes motions in inertial space
for mission objectives other than momentum management. As an
advantageous result, using the presently disclosed techniques, the
torque authority provided by a single gimbaled thruster may, over
time, be used to span a substantial portion of the entire 3D
inertial space.
[0060] As a further example, the present techniques may be useful
for a small spacecraft orbiting the Earth in a low earth orbit,
having control moment gyros (CMGs). The purpose of this spacecraft
may be reconnaissance, whereby it periodically takes images of the
Earth, by rapidly slewing its main bus using the CMGs. For this
type of a satellite, the momentum is typically accumulated in a
fixed direction in inertial space. If a single thruster aimed
through the center of mass is present, than the entire satellite
may be rapidly steered using the CMGs in 3D space, to quickly dump
full 3-axis momentum.
[0061] Referring now to FIG. 8, method 800 for managing momentum of
a spacecraft traveling along a trajectory will be described. As the
spacecraft travels along the trajectory, a desired line of thrust
undergoes a substantial rotation in inertial space. At block 810, a
respective momentum storage error (MSE) may be determined for each
of three inertial spacecraft axes. The respective MSE may represent
a difference, for each respective axis, between a momentum value
actually stored and a desired momentum value.
[0062] At block 820, each respective MSE may be reduced by
producing a plurality of torques with at least one thruster
disposed on a single positioning mechanism. More particularly, at
block 821, a first torque may be produced by orienting the
thruster, with the single positioning mechanism. The first torque
may reduce the respective MSE of at least one of a first and a
second of the three inertial spacecraft axes when the spacecraft is
located at a first location on the trajectory. Subsequently, at
block 822, when the spacecraft is located at a second location on
the trajectory, a second torque may be produced that reduces the
respective MSE of at least a third of the three inertial spacecraft
axes by orienting the thruster, with the single positioning
mechanism.
[0063] Advantageously, spacecraft acceleration and MSE may be
simultaneously controlled by the at least one thruster disposed on
the single positioning mechanism.
[0064] Referring now to FIG. 9, method 900 will be described. At
block 910, an optimal satellite steering profile for a spacecraft
may be dynamically computed, based on a position of the spacecraft
on a trajectory. The spacecraft may include at least one thruster
disposed on a single positioning mechanism, an orbit propagator or
GPS, and an inertial reference sensor. For example, the thruster
may be a chemical thruster, or low thrust electric propulsion
device such as a Hall effect thruster, or SPT. The inertial
reference sensor may be an integrating rate gyro, or star tracker,
for example.
[0065] At block 920, the satellite's actual position may be
dynamically computed using, for example, outputs from the orbit
propagator or GPS.
[0066] At block 930, the satellite may be steered according to the
computed optimal satellite steering profile. For example, the
thruster may be used to impart a change in velocity of the
satellite along a desired direction;
[0067] At block 940, the thruster may be periodically shut down,
and the inertial reference sensor may be re-initialized. For
example, the inertial reference sensor may be re-initialized with
respect to a reacquired celestial body, such as the Earth. As a
further example, where the inertial references sensor is a gyro, it
may be reset to remove any drift when the spacecraft is in an Earth
pointed orientation, using a calculated position of the Earth
relative to the satellite, satellite orbital information and Earth
sensor data.
[0068] At block 950, the thruster may be restarted.
[0069] As desired, blocks 910 through 950 may be autonomously
repeated until the desired orbit is reached. During such time,
three axis momentum management of the satellite may be performed by
the method illustrated in FIG. 8.
[0070] Thus, spacecraft momentum management techniques have been
disclosed, whereby, for a spacecraft traveling along a trajectory,
three-axis spacecraft momentum management may be performed by an
actuator consisting of as few as one thruster disposed on a single
positioning mechanism. It will be understood that, although a
benefit of the present teachings is to enable three-axis spacecraft
momentum management to be performed as few as one thruster disposed
on a single positioning mechanism, for redundancy purposes, two or
more positioning mechanisms may be provided without departing from
the scope of the present teachings.
[0071] The foregoing merely illustrates principles of the
invention. It will thus be appreciated that those skilled in the
art will be able to devise numerous systems and methods which,
although not explicitly shown or described herein, embody said
principles of the invention and are thus within the spirit and
scope of the invention as defined by the following claims.
* * * * *